U.S. patent number 8,356,978 [Application Number 12/623,666] was granted by the patent office on 2013-01-22 for turbine airfoil platform cooling core.
This patent grant is currently assigned to United Technologies Corporation. The grantee listed for this patent is Jeffrey S. Beattie, Matthew A. Devore, Matthew S. Gleiner, Douglas C. Jenne. Invention is credited to Jeffrey S. Beattie, Matthew A. Devore, Matthew S. Gleiner, Douglas C. Jenne.
United States Patent |
8,356,978 |
Beattie , et al. |
January 22, 2013 |
Turbine airfoil platform cooling core
Abstract
A gas turbine engine component has a platform and an airfoil
extending from the platform. The platform has a pressure side and a
suction side. A cooling passage is formed within the platform, and
extends along a pressure side of the platform. Air leaves the
passage through an air outlet on a suction side of the
platform.
Inventors: |
Beattie; Jeffrey S. (South
Glastonbury, CT), Devore; Matthew A. (Manchester, CT),
Gleiner; Matthew S. (Vernon, CT), Jenne; Douglas C.
(West Hartford, CT) |
Applicant: |
Name |
City |
State |
Country |
Type |
Beattie; Jeffrey S.
Devore; Matthew A.
Gleiner; Matthew S.
Jenne; Douglas C. |
South Glastonbury
Manchester
Vernon
West Hartford |
CT
CT
CT
CT |
US
US
US
US |
|
|
Assignee: |
United Technologies Corporation
(Hartford, CT)
|
Family
ID: |
43611936 |
Appl.
No.: |
12/623,666 |
Filed: |
November 23, 2009 |
Prior Publication Data
|
|
|
|
Document
Identifier |
Publication Date |
|
US 20110123310 A1 |
May 26, 2011 |
|
Current U.S.
Class: |
416/193A;
415/115; 416/97R |
Current CPC
Class: |
B22C
9/10 (20130101); F01D 9/02 (20130101); F01D
5/18 (20130101); F05D 2260/20 (20130101); F05D
2240/81 (20130101) |
Current International
Class: |
F01D
5/00 (20060101) |
Field of
Search: |
;415/115,116
;416/90R,92,97R,193A |
References Cited
[Referenced By]
U.S. Patent Documents
Primary Examiner: Nguyen; Ninh H
Assistant Examiner: McDowell; Liam
Attorney, Agent or Firm: Carlson, Gaskey & Olds, PC
Government Interests
This invention was made with government support under Contract No.
F33615-03-D-2354-0009 awarded by the United States Air Force. The
Government may therefore have certain rights in this invention.
Claims
What is claimed is:
1. A gas turbine engine component comprising: a platform, and an
airfoil extending from said platform, said platform having a
pressure side and a suction side; a cooling passage located within
said platform, and extending along the pressure side of said
platform, and an outlet for air leaving said cooling passage, said
outlet being on the suction side of said platform; and said outlet
is at a radially outer face of said platform, and not through an
edge of said platform.
2. The component as set forth in claim 1, wherein an extension of a
trailing edge of said airfoil can be extended to a point on a side
wall of said platform, and said cooling passage is on one side of
said point, and said outlet being on an opposed side.
3. The component as set forth in claim 1, wherein said cooling
passage passes beneath a portion of said airfoil between an inlet
and said outlet.
4. The component as set forth in claim 3, wherein said cooling
passage passes beneath a trailing edge of said airfoil, and to said
suction side.
5. The component as set forth in claim 3, wherein said airfoil has
internal cooling passages, and said cooling passage passes beneath
one of said internal cooling passages in said airfoil before
reaching said outlet on said suction side.
6. The component as set forth in claim 1, wherein said cooling
passage does not pass underneath said airfoil, but instead is
positioned between a trailing edge of said airfoil, and a side wall
of said platform when passing from said pressure side to said
suction side.
7. The component as set forth in claim 1, wherein an end of said
cooling passage, and on said suction side, leading to said outlet
curves toward a first side wall of said platform, and then turns
back to an opposed side wall of said platform.
8. The component as set forth in claim 1, wherein said cooling
passage has a bulged intermediate portion to increase heat transfer
by increasing contact area between said cooling passage and a
portion of said platform.
9. The component as set forth in claim 1, wherein cooling
structures elements are positioned within said cooling passage.
10. The component as set forth in claim 1, wherein said component
is a turbine blade.
11. The component as set forth in claim 1, wherein said component
is a static vane.
12. The component as set forth in claim 11, wherein said static
vane has a platform at both a radially outer edge and a radially
inner edge.
13. The component as set forth in claim 12, wherein said cooling
passage is located in said radially outer edge platform.
14. The component as set forth in claim 1, wherein said radially
outer face faces radially inwardly.
15. The component as set forth in claim 1, wherein said radially
outer face faces radially outwardly.
16. A gas turbine engine component comprising: a platform, and an
airfoil extending from said platform, said platform having a
pressure side and a suction side; a cooling passage formed within
said platform, and extending along a pressure side of said
platform, and an outlet for air leaving said passage, said outlet
being on a suction side of said platform; an extension of a
trailing edge of said airfoil can be extended to a point on said
side wall of said platform, and said inlet to said cooling passage
will be on one side of said point, and said outlet being on an
opposed side; and said outlet is at a radially outer face of said
platform.
17. The gas turbine engine as set forth in claim 16, wherein said
radially outer face faces radially inwardly.
18. The gas turbine engine as set forth in claim 16, wherein said
radially outer face faces radially outwardly.
Description
BACKGROUND OF THE INVENTION
This application relates to a cooling passage for a platform in a
gas turbine component.
Gas turbine engines include a compressor which compresses air and
delivers it downstream into a combustion section. The air is mixed
with fuel in the combustion section and ignited. Products of this
combustion pass downstream over turbine rotors, which are driven to
rotate. In addition, static vanes are positioned adjacent to the
turbine rotors to control the flow of the products of
combustion.
The turbine rotors carry blades. The blades and the static vanes
have airfoils extending from platforms. The blades and vanes are
subject to extreme heat, and thus cooling schemes are utilized for
each.
It is known to provide a cooling passage in the platform of the
vanes and blades to cool the platform on the pressure side. Such
passages have an outlet on the pressure side of the platform.
SUMMARY OF THE INVENTION
A gas turbine engine component has a platform and an airfoil
extending from the platform. The platform has a pressure side and a
suction side. A cooling passage is located within the platform, and
extends along a pressure side of the platform. Air leaves the
passage through an air outlet on a suction side of the
platform.
These and other features of the present invention can be best
understood from the following specification and drawings, the
following of which is a brief description.
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1 shows a turbine rotor.
FIG. 2 is a partial view of a turbine blade.
FIG. 3 is a cross-sectional view through the platform of the FIG. 2
blade.
FIG. 4 is a top view of a first embodiment.
FIG. 5 shows a second embodiment.
FIG. 6A shows yet another embodiment.
FIG. 6B shows a portion of the FIG. 6A embodiment.
FIG. 7 shows a static vane.
FIG. 8 is a top view of the FIG. 7 vane.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT
FIG. 1 shows a turbine section 20 including a rotor 22 carrying a
blade 24. Blade 24 includes a platform 28 and an airfoil 30. As
also shown, a vane 11 is positioned adjacent to the blade 24.
As shown in FIG. 2, airfoil 30 has a leading edge 31 and a trailing
edge 33. A pressure side 32 of the airfoil is shown in this Figure.
A cooling passage 34 is positioned on the pressure side of the
airfoil, and in the platform 28. The cooling passage 34 extends to
an outlet 40, which, as will be explained below, sits on a suction
side of the platform 28. The blade 24 includes a root section 26
which is utilized to secure the blade within the rotor. In
addition, a plurality of cooling passages 36 and 38 extend through
the root 26 from a cooling air supply and upwardly into the airfoil
30, as known.
As shown in FIG. 3, the cooling passage 34 has an inlet 42 for
supplying air. As shown, the inlet 42 comes into the platform 28 at
a lower surface, and rearward of a leading edge 100 of the platform
28. Cooling air passes into an inlet 42, through the cooling
passage 34, and outwardly of the outlet 40 cooling the platform 28.
The inlet 42 to the cooling passage 34 can be from any number of
locations depending on the particular design, and the environment
in which the component is to be utilized. A worker of ordinary
skill in the art would be able to identify any number of potential
sources of cooling air. As shown, a source of air communicates to
the inlet.
As can be appreciated from FIG. 4, the airfoil 30 has a suction
side 50. The outlet 40 of the cooling passage 34 is on the suction
side of the platform. Stated another way, should the airfoil be
extended from the trailing edge 33 to the edge 103 of the platform
28, it will be at a position X. This could be defined as a dividing
line between the pressure and suction sides of the platform. The
outlet 40 is on the suction side.
In the FIG. 4 embodiment, the cooling passage 34 passes through the
platform, and beneath the trailing edge 33 before getting to the
outlet 40. As can be appreciated also from this Figure, the end 102
of the cooling passage curves away from the edge 103, before
curving back toward the edge 103 and reaching outlet 40. The curve
shown at the end 102, and leading toward the outlet 40, assists in
directing the exiting air flow to line up with the main gas air
flow through the gas turbine engine. However, a straight passage to
the outlet may also be utilized. As shown, the cooling passage has
a bulged intermediate portion 400. The bulged portion 400 increases
the cooling surface area at a particular location along the path,
and further allows better heat transfer characteristics.
Various cooling structures may be included in the cooling passage
34. Pin fins, trip strips, guide vanes, pedestals, etc., may be
placed within the passage to manage stress, gas flow, and heat
transfer. As shown, a number of pins 21 may be formed within the
cooling passage 34 to increase the heat transfer effect. As
mentioned, any number of other heat transfer shapes can be
utilized, including a rib 52 adjacent the outlet. Further, if there
are localized hot spots, outlet holes can be formed either to the
outer face of the platform, or to the outer edge 103, as deemed
appropriate by the designer. Additionally, holes can be drilled
from the underside of the platform to supply additional air to the
passage.
As is clear, the curving ends 102 and 150 are located on the
suction sides of their respective embodiments.
As shown in FIG. 5, a second embodiment 124 has platform 128, and
platform cooling passage 134. Again, an extension from the trailing
edge 133 of the airfoil 130 reaches point X. The cooling passage
134 passes around the airfoil trailing edge 133, and the outlet 152
of the cooling passage 134 is on the suction side of point X, and
the suction side of the platform 128. Stated another way, the
cooling passage does not pass underneath the airfoil, but instead
is positioned between the trailing edge 133 and the side wall of
the platform when passing from the pressure side to the suction
side. Again, the end 150 curves away from the edge 103, and a rib
151 is included.
All of the above discussed cooling features, such as features 136
and 151, and holes can be utilized.
FIG. 6A shows yet another embodiment 160 having a platform 165, and
an airfoil 162. Here, the cooling passage 166 has a serpentine
path, including a curve 168 on the pressure side, which leads to a
leading edge extending portion 170, a crossing portion 172, a
portion 174, which is now on the suction side, and which leads to a
final portion 176 leading to the outlet 178. Again, the outlet 178
is on the suction side, and on an opposed side of the point X from
the inlet to the cooling passage 166.
In the FIG. 6A embodiment, a central passage 164 in the airfoil 162
can be seen to have the cooling passage portion 172 passing
underneath.
As shown in FIG. 6B, the passage 172 preferably does not
communicate with the passage 164 when passing underneath the
passage 164. In addition, while the serpentine passage 166 is
disclosed, a more direct route underneath the airfoil can also be
utilized.
The inlet to the cooling passages in FIGS. 4-6 may be positioned
anywhere, as mentioned above.
An embodiment 200 is shown in FIG. 7, wherein the cooling passage
is incorporated into a static vane arrangement. As shown, vane
airfoils 208 and 206 extend between platforms 202 and 204. The
platform 204 will be a radially inner end wall when the vane
embodiment 200 is mounted within an engine, while the platform 202
will be radially outwardly. While a dual vane arrangement is shown,
a single vane may also incorporate the cooling passage, as may any
number of other static vane arrangements.
As shown in FIG. 8, again, a cooling passage 212 is formed on a
pressure side 210 of the airfoil 208. The outlet 214 is again on
the suction side 211, and on an opposed side of the point X from
the inlet to the cooling passage 212.
As can be appreciated from the several embodiments, the outlet is
located on a radially outer face of the platforms, and not through
the edge 103. The above is true of all of the embodiments. In the
vane embodiments, the "outer face" is facing radially inwardly, but
from a functional standpoint, the face of the platform from which
the airfoil extends is the "radially outer face" for purposes of
this application.
The cooling passages 34 may be formed from any suitable core
material known in the art. For example, the cooling passage 34 may
be formed from a refractory metal or metal alloy such as molybdenum
or a molybdenum alloy. Alternatively, the cooling passage 34 may be
formed from a ceramic or silica material.
The cooling passage 34 can be formed by a lost core molding
technique, as is known in the art. Alternatively, the passage can
be created by welding a plate onto the part after the passage has
been created by a molding technique. Any number of other ways of
forming such internal structure can also be utilized.
The platform cooling passage provides shielding to the
underplatform from hot gases. Shielding reduces heat pick-up in the
rim, potentially improving rotor/seal/damper, etc. life. Shielding
also reduces bulk panel temperatures, which increases creep life on
the end wall.
Although several embodiment of this invention have been disclosed,
a worker of ordinary skill in this art would recognize that certain
modifications would come within the scope of this invention. For
that reason, the following claims should be studied to determine
the true scope and content of this invention.
* * * * *